The present invention relates generally to spacecraft architecture, and more specifically, to a spacecraft control architecture that provides a payload environment that is free from spacecraft-borne vibrations while still being able to control the motion of the payload in space.
In many spacecraft-borne missions, it is imperative to attenuate mechanical disturbances and thus isolate a payload from vibrations generated on the spacecraft while retaining the capability of precisely controlling the motion of the payload. The problem caused by vibrations is common to scientific, commercial and military missions. Examples of scientific missions that are particularly affected by spacecraft vibrations are space-based telescopes, such as the Hubble Space Telescope and the proposed Next Generation Space Telescope, and space-based interferometers, such as NASA""s Space Interferometery Mission. In the future, the use of laser-based communications, requiring precision pointing between satellites, will likely increase the significance of the problem for commercial payloads. Finally, motion stability and precision pointing and tracking are key technologies for military apparatus such as earth observatories and space-based defense and missile systems, such as the proposed space-based laser systems.
As reported in NASA Technical Memorandum 106496, titled Final Reportxe2x80x94Vibration Isolation Technology (VIT) A TD Project by J. Lubomski et al. (March 1994), a wide variety of vibration isolation technologies have been considered. Passive isolation tends to be more cost effective, but has limited effectiveness specially at low frequencies. In general, active systems require sensing of motion or position, and a feedback and/or feedforward control loop to counteract mechanical excitation and minimize motion of an isolated body. Such systems typically introduce the complexity of a high-gain control system and isolation performance is limited by sensor characteristics.
Z. Geng et al. describes a vibration isolation system for space-borne structures in the Journal of Intelligent Material Systems and Structures, Vol. 6 (November 1995). The system includes an apparatus for providing real-time, six degree-of-freedom active vibration isolation. The apparatus requires multiple accelerometers and at least six force sensor inputs, six analog outputs, a sixteen channel digital I/O, and extensive computation power to accommodate complex control algorithms. The active system of Z. Geng et al. includes two layers of six vibration control mechanisms. An upper layer is the six degrees-of-freedom active vibration isolator with a mobile plate and base plate connected by six active elements. Each active element consists of a Terfenol-D actuator, a force sensor, a pair of accelerometers, and a pair of flexible joints. Six accelerometers are mounted on the mobile plate to measure the residual vibration and another six are placed on the base plate to monitor base plate excitation. Both acceleration and force measurements are fed to signal conditioners and then delivered to a control system, which generates control signals to power amplifiers that drive the Terfenol-D actuators. This is a complex system with performance limited both by the mechanical connection between the two bodies and by the characteristics of the sensors used to measure residual vibrations.
U.S. Pat. No. 4,848,525 to Jacot et al. discloses a dual mode vibration isolator for actively isolating vibrations between a forward body and an aft body. The isolator includes an intermediate mechanical stage referred to as a xe2x80x9cmounting memberxe2x80x9d. The mounting member is pivotally connected via three pairs of linear actuators to the aft body. The mounting member exerts forces on the forward body via the use of three pairs of magnetic actuators. The linear actuators extend and contract to reposition the mounting member on the aft body, which in turn repositions the forward body relative to the aft body. Each magnetic actuator has an accompanying flux sensor and gap sensor, and each linear actuator is paired with a length sensor. Except for the additional mechanical stage, used to extend the range of motion between the forward and aft bodies at the expense of significant additional complexity, this is a typical magnetic isolation system with limited performance at low frequencies.
Edberg et. al. (AAS-96-071) describe the STABLE micro-gravity isolation system which exemplifies the state-of-the-art in magnetic isolation systems. The isolated payload is levitated by three dual-axis wide gap electromagnetic actuators. Signals from accelerometers located on the payload are used by a high bandwidth feedback controller to command counteracting electromagnetic actuator forces. Signals from three, two-axis optical sensors measure the position of the payload with respect to the mounting base and are used in a low bandwidth position loop to command the electromagnetic actuators. to keep the payload centered with respect to the base. As indicated by Edberg et. al., performance is limited by sensor characteristics, such as accelerometer noise floor. In addition, the control loop architecture limits isolation performance at low frequencies, near the bandwidth of the position control loop.
The problems with these and other existing vibration isolation systems are performance limitations at low frequencies, which is inherent to the architecture of these systems and specifically their control system architecture, and limitations associated with load, position, velocity and acceleration sensing, which, on existing systems, directly affect isolation performance.
What is needed is a system that exhibits superior vibration isolation performance down to very low frequencies, and isolation performance that is not limited by sensor characteristics. What is also needed is a system that allows control of the motion of the payload without limitation on its range of motion while using a small number of sensors and actuators to avoid the low reliability and high costs of a complex system.
An object of the present invention is to provide a spacecraft architecture and in particular a spacecraft control architecture that simultaneously addresses the problems of payload pointing and motion control and vibration isolation, while providing superior vibration isolation down to zero frequency.
Another object of this invention is to provide a control architecture that allows the isolation of a payload from spacecraft-borne vibrations down to zero frequency, and a system in which the isolation performance is not limited by sensor characteristics.
Another object of this invention is to provide control of the attitude of a payload to any desired orientation without limitation on its range of motion, while isolating the payload from spacecraft-borne vibrations.
Still another object of the present invention is to provide precision pointing and motion control and vibration isolation for applications requiring high levels of motion control and stability, such as imaging payloads, laser-based communications and tracking systems.
The present invention achieves these and other objects by providing a system that includes a payload module and a support module that are preferably mechanically de-coupled. The motion of the payload module is controlled by reacting on the support module using non-contact actuators disposed between the two modules. The motion of the support module is controlled to follow the payload module using external actuators that react against the surroundings. In this way, no forces are applied between the payload and support modules due to relative motion control and vibration isolation is achieved down to zero frequency. Moreover, vibration isolation is not limited by sensor characteristics. In fact, if the sensors used to measure the motion of the payload module with respect to its surroundings stopped functioning, the payload module would be a drift, but the support module would continue to follow the payload module and the vibration isolation performance would be unaffected. In the system of this invention, the payload and support modules fly in formation in close-proximity and interact through non-contact actuators to achieve precision motion control and high motion stability of the payload module.
Specifically, the system of the present invention includes a spacecraft comprising of a payload module and a support module that are preferably mechanically de-coupled, and non-contacting means of measuring relative position and applying forces between the two modules. The payload module contains critical components requiring precision motion control and high motion stability, e.g. a telescope, a communications or a tracking system. The support module contains mission support equipment that does not require precise motion control and high motion stability. It also contains the main sources of vibration, such as reaction wheels and thrusters, and large flexible appendages, such as solar panels and sun-shields. Non-contact position sensors, disposed between the payload and support modules, are used to obtain information on the relative position and attitude between the payload and support modules: These sensors can be based on various technologies for non-contact measurement of distance, such as inductive, capacitive, or optical. Non-contact actuators, also disposed between the payload and support modules, are used to apply forces between the payload and support modules and control the motion of the payload module. These actuators can be electromagnetic, such as voice-coil actuators.
In a preferred embodiment, six non-contact position sensors and six non-contact actuators are arranged between the payload and support modules in a hexapod configuration. The non-contact actuators are voice-coil actuators with the field assembly preferably mounted on the payload module and the coil assembly mounted on the support module. Force control on each non-contact actuator allows relative motion between the payload and support modules without transmission of vibrations between the two modules. Force control is achieved with a high-bandwidth current control loop on the coil assembly of each voice-coil actuator which counter-act vibration induced currents on the coil assembly
The system of the present invention also includes external actuators located on the support module, i.e. actuators that react against the surroundings, and payload or target sensors, i.e. sensors that provide information on the payload position and attitude with respect to its surroundings. The external actuators are used to move the support module by reacting against its surroundings, e.g. inertial space, and follow the motion of the payload module. Examples of external actuators for space application are thrusters, reaction wheels, magnetic torquers, control moment gyros, and solar sails. The payload or target sensors provide information required to control the motion of the payload with respect to its surroundings, e.g. inertial space. Examples of payload or target sensors for space applications are star-trackers, accelerometers and gyroscopes.
According to the control architecture of the present invention, a payload position and attitude control unit receives the signals from the payload or target sensors and outputs a command to the non-contact actuators that generate a force and/or moment between the payload and support modules to move the payload module to a desired position and/or orientation. At the same time, a relative position and attitude control unit receives the signals from the relative position sensors and outputs commands to the external actuators that move the support module by reacting against its surroundings, so that the support module follows the motion of the payload module. The external actuators thus move the support module to maintain a desired relative position and attitude with respect to the payload module. While the motion of the payload module is controlled with a high-degree of precision and is very stable, the motion requirements for the support module are much more relaxed. The support module merely must stay within a sufficient distance and angular range of the payload module in order to prevent the non-contact position sensors and non-contact actuators from going out of range.
A key element of the control architecture of this invention is that no forces are applied between the payload and support modules to maintain relative motion control, this allows vibration isolation down to zero frequency. Another important advantage of this architecture. is that the performance of the isolation system is not dependent on sensor characteristics and therefore, superior isolation can be achieved. The system and control architecture of the present invention also allows the payload to be pointed over the entire celestial sphere (4xcfx80 steradian).
In applications where some mechanical coupling is required between the payload and support modules, for example due to data and power cables and/or cooling lines, the non-contact position sensors and non-contact actuators can be used in closed-loop control to cancel the effect of the mechanical coupling between the two modules. This represents a simple modification in the control logic and does not require any additional hardware. In practice, the cancellation is not perfect and the isolation performance is reduced when compared with an equivalent mechanically de-coupled system.
A variation of the described control architecture, which is also part of the present invention, uses the non-contact actuators to perform relative position (translation) control and external actuators to perform relative attitude control between the payload and support modules. In this case, the forces on the non-contact actuators for relative position (translation) control are computed to generate zero moment about the center of mass of the payload module and therefore do not affect the attitude of the payload module. This variation is of particular interest for space applications where, in general, precise attitude control is required while control of translational motion is less important. In addition, in such applications the use of external actuators to maintain relative position (translation) control may, in some cases, present difficulties. For example, the use of thrusters on long duration missions may require a large amount of propellant that is not feasible to carry on-board of the spacecraft.
These and other objects and features of the invention will be better understood in referencing the following detailed description and drawings.